US20160245091A1 - Gas turbine engine airfoil with auxiliary flow channel - Google Patents
Gas turbine engine airfoil with auxiliary flow channel Download PDFInfo
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- US20160245091A1 US20160245091A1 US15/029,802 US201415029802A US2016245091A1 US 20160245091 A1 US20160245091 A1 US 20160245091A1 US 201415029802 A US201415029802 A US 201415029802A US 2016245091 A1 US2016245091 A1 US 2016245091A1
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- flow channel
- airfoil
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- gas turbine
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/141—Shape, i.e. outer, aerodynamic form
- F01D5/145—Means for influencing boundary layers or secondary circulations
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow blades
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D17/00—Regulating or controlling by varying flow
- F01D17/10—Final actuators
- F01D17/12—Final actuators arranged in stator parts
- F01D17/14—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D17/00—Regulating or controlling by varying flow
- F01D17/10—Final actuators
- F01D17/12—Final actuators arranged in stator parts
- F01D17/14—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits
- F01D17/16—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes
- F01D17/162—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes for axial flow, i.e. the vanes turning around axes which are essentially perpendicular to the rotor centre line
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/141—Shape, i.e. outer, aerodynamic form
- F01D5/146—Shape, i.e. outer, aerodynamic form of blades with tandem configuration, split blades or slotted blades
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
- F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
- F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/40—Casings; Connections of working fluid
- F04D29/52—Casings; Connections of working fluid for axial pumps
- F04D29/54—Fluid-guiding means, e.g. diffusers
- F04D29/56—Fluid-guiding means, e.g. diffusers adjustable
- F04D29/563—Fluid-guiding means, e.g. diffusers adjustable specially adapted for elastic fluid pumps
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/66—Combating cavitation, whirls, noise, vibration or the like; Balancing
- F04D29/68—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers
- F04D29/681—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers especially adapted for elastic fluid pumps
- F04D29/682—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers especially adapted for elastic fluid pumps by fluid extraction
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/66—Combating cavitation, whirls, noise, vibration or the like; Balancing
- F04D29/68—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers
- F04D29/681—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers especially adapted for elastic fluid pumps
- F04D29/684—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers especially adapted for elastic fluid pumps by fluid injection
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2220/00—Application
- F05D2220/30—Application in turbines
- F05D2220/32—Application in turbines in gas turbines
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/10—Stators
- F05D2240/12—Fluid guiding means, e.g. vanes
- F05D2240/123—Fluid guiding means, e.g. vanes related to the pressure side of a stator vane
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/10—Stators
- F05D2240/12—Fluid guiding means, e.g. vanes
- F05D2240/124—Fluid guiding means, e.g. vanes related to the suction side of a stator vane
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/20—Rotors
- F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
- F05D2240/305—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the pressure side of a rotor blade
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/20—Rotors
- F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
- F05D2240/306—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the suction side of a rotor blade
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/35—Combustors or associated equipment
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/60—Fluid transfer
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- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Definitions
- This disclosure relates to a gas turbine engine, and more particularly to a gas turbine engine airfoil having an auxiliary flow channel for receiving and communicating a portion of core airflow through the airfoil.
- Gas turbine engines typically include at least a compressor section, a combustor section, and a turbine section.
- air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases.
- the hot combustion gases flow through the turbine section, which extracts energy from the hot combustion gases to power the compressor section as well as other gas turbine engine loads.
- the compressor and turbine sections typically include alternating rows of rotating blades and flow directing vanes.
- the rotating blades extract energy from the core airflow that is communicated through the gas turbine engine, while the vanes direct the core airflow to a downstream row of blades.
- the vanes can be manufactured to a fixed flow area that is optimized for a single flight point.
- a component for a gas turbine engine includes, among other things, an airfoil that includes a pressure side surface and a suction side surface that join together at a leading edge and a trailing edge and a flow channel that extends between the pressure side surface and the suction side surface.
- the component is a vane or a blade.
- the component is a mid-turbine frame, an exit guide vane or a fan blade.
- the flow channel includes an inlet at the pressure side surface and an outlet at the suction side surface.
- the flow channel includes a radial dimension that is less than a total span of the airfoil.
- the flow channel includes an outlet positioned upstream from a throat area that extends between the airfoil and an adjacent airfoil.
- the flow channel includes an outlet positioned downstream from a throat area that extends between the airfoil and an adjacent airfoil.
- the flow channel defines a flow passage that extends from a tip toward a root of the airfoil.
- the flow channel defines a flow passage that extends from a root toward a tip of the airfoil.
- the flow channel extends along a curved path.
- a gas turbine engine includes, among other things, a variable area section, a variable vane disposed within the variable area section and a blade downstream from the variable vane. At least one of the variable vane and the blade includes a flow channel having an inlet at a pressure side surface and an outlet at a suction side surface of an airfoil of either the variable vane or the blade.
- variable area section is a variable area turbine section.
- variable area section is a variable area compressor section.
- both the variable vane and the blade include the flow channel.
- the flow channel extends in each of a chordwise and a spanwise direction of the airfoil.
- a method of operating a gas turbine engine includes, among other things, communicating a portion of core airflow through a flow channel that extends between a pressure side surface and a suction side surface of at least one airfoil of the gas turbine engine.
- the method includes the step of rotating a variable vane prior to the step of communicating.
- the method includes communicating the portion of airflow into an inlet of the flow channel located along the pressure side surface and expelling the portion of airflow through an outlet of the flow channel located along the suction side surface.
- the method includes the step of influencing incidence angle variation of a downstream component with the portion of core airflow.
- the method includes the step of communicating a second portion of core airflow through a flow channel of the downstream component.
- FIG. 1 illustrates a schematic, cross-sectional view of a gas turbine engine.
- FIG. 2 illustrates a variable area section of a gas turbine engine.
- FIG. 3 illustrates an airfoil having a flow channel.
- FIG. 4 illustrates another airfoil flow channel.
- FIG. 5 illustrates another airfoil flow channel.
- FIG. 6 illustrates yet another airfoil flow channel.
- FIGS. 7A and 7B illustrate vanes having flow channels.
- FIGS. 8A and 8B illustrate blades having flow channels.
- FIGS. 9A, 9B and 9C illustrate multiple configurations of a variable area gas turbine engine section that includes one or more components having flow channels.
- a variable area section of a gas turbine engine may employ one or more airfoils that include a flow channel extending between a pressure side surface and a suction side surface.
- Core airflow may be communicated through the flow channel to influence a location of the flow stagnation point on the airfoil and downstream components.
- a portion of core airflow may be communicated through the flow channel in order to modify the incidence angle of the core flow on the airfoil.
- a portion of core airflow may be communicated through the flow channel in order to modify the incidence angle at which core airflow impinges on downstream hardware.
- FIG. 1 schematically illustrates a gas turbine engine 20 .
- the gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22 , a compressor section 24 , a combustor section 26 and a turbine section 28 .
- Alternative engines might include an augmentor section (not shown) among other systems or features.
- the fan section 22 drives air along a bypass flow path B in a bypass duct defined within a nacelle 15
- the compressor section 24 drives air along a core flow path C for compression and communication into the combustor section 26 then expansion through the turbine section 28 .
- the exemplary engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38 . It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of bearing systems 38 may be varied as appropriate to the application.
- the low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42 , a first (or low) pressure compressor 44 and a first (or low) pressure turbine 46 .
- the inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30 .
- the high speed spool 32 includes an outer shaft 50 that interconnects a second (or high) pressure compressor 52 and a second (or high) pressure turbine 54 .
- a combustor 56 is arranged in exemplary gas turbine 20 between the high pressure compressor 52 and the high pressure turbine 54 .
- a mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46 .
- the mid-turbine frame 57 further supports bearing systems 38 in the turbine section 28 .
- the inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.
- the core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52 , mixed and burned with fuel in the combustor 56 , then expanded over the high pressure turbine 54 and low pressure turbine 46 .
- the mid-turbine frame 57 includes airfoils 59 which are in the core airflow path C.
- the turbines 46 , 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion.
- gear system 48 may be located aft of combustor section 26 or even aft of turbine section 28
- fan section 22 may be positioned forward or aft of the location of gear system 48 .
- the engine 20 in one example is a high-bypass geared aircraft engine.
- the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10)
- the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3
- the low pressure turbine 46 has a pressure ratio that is greater than about five.
- the engine 20 bypass ratio is greater than about ten (10:1)
- the fan diameter is significantly larger than that of the low pressure compressor 44
- the low pressure turbine 46 has a pressure ratio that is greater than about five 5:1.
- Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle.
- the geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.
- the fan section 22 of the engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet.
- TSFC Thrust Specific Fuel Consumption
- Low fan pressure ratio is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system.
- the low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45.
- Low corrected fan tip speed is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)] 0.5 .
- the “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second.
- FIG. 2 illustrates a variable area section 100 that may be incorporated into a gas turbine engine, such as the gas turbine engine 20 of FIG. 1 .
- the variable area section 100 is a variable area turbine section.
- the variable area section 100 is a variable area compressor section.
- the variable area section 100 may make up part of or the entirety of the compressor section 24 and/or the turbine section 28 of the gas turbine engine 20 . In other words, these sections could have components in addition to what is shown in FIG. 2 .
- variable area section 100 includes one or more stages of alternating rows of vanes and blades.
- the variable area section 100 includes a single stage (i.e., one row of vanes and one row of blades); however, additional stages could be incorporated into the variable area section 100 .
- a case structure 62 houses the components of the variable area section 100 .
- the case structure 62 is an outer engine casing that circumscribes the components of the variable area section 100 .
- variable area section 100 includes a variable vane assembly 64 having at least one variable vane 66 .
- the variable vane assembly 64 could include an array of variable vanes 66 circumferentially disposed about the engine centerline longitudinal axis A.
- the variable vane assembly 64 could include a combination of both fixed and variable vanes.
- variable vanes 66 are selectively configurable to change a flow parameter associated with the variable area section 100 .
- each variable vane 66 may be rotated or pivoted (via an actuation system) about a spindle axis SA in order to change the rotational positioning of the variable vane 66 .
- Rotating the variable vane(s) 66 changes the amount of core airflow F that enters the variable area section 100 , thereby influencing the flow area of the variable area section 100 .
- variable vane(s) 66 is rotatable relative to an inner platform 68 and an outer platform 70 of the variable vane assembly 64 .
- the inner platform 68 and the outer platform 70 may be mounted to the case structure 62 in any known manner.
- a rotor assembly 72 is positioned downstream from the variable vane assembly 64 .
- the rotor assembly 72 includes at least one rotor disk 74 that carries one or more rotor blades 76 .
- the rotor blades 76 rotate about the engine central longitudinal axis A to extract energy from the core airflow F, thereby moving the rotor disk 74 and powering various gas turbine engine loads.
- the rotor blades 76 rotate relative to blade outer air seals (BOAS) 78 that establish a radially outer flow path boundary for channeling the core airflow F through the variable area section 100 .
- the BOAS 78 may mount to the case structure 62 and extend in relationship to a tip of each rotating blade 76 in order to seal between the blades 76 and the case structure 62 .
- variable vane(s) 66 can change the incidence angle ⁇ at which core airflow F impinges upon the rotor blades 76 of the downstream rotor assembly 72 .
- Incidence angle variation can negatively influence gas turbine engine efficiency by altering the stagnation point on the airfoil.
- the variable vanes 66 and/or rotor blades 76 may include airfoils having auxiliary flow channels for addressing such variations. Airfoils of this type are discussed in greater detail below with respect to FIGS. 3-9 .
- FIG. 3 illustrates a cross-sectional view of an exemplary airfoil 80 that includes a flow channel 82 .
- the airfoil 80 may be part of a vane and/or a blade.
- the airfoil 80 could be a section of a variable vane 66 of the variable vane assembly 64 and/or a rotor blade 76 of the rotor assembly 72 of FIG. 2 (see, for example, FIGS. 7A, 7B and 8A, 8B ).
- the airfoil 80 could also be part of other gas turbine engine components including but not limited to the mid-turbine frame, a fan exit guide vane or a fan blade.
- the flow channel 82 extends inside of the airfoil 80 between a pressure side surface 84 and a suction side surface 86 of the airfoil 80 .
- the pressure side surface 84 and the suction side surface 86 are spaced apart from one another and generally join together at a leading edge 88 and a trailing edge 90 of the airfoil 80 .
- the flow channel 82 includes an inlet 92 located at the pressure side surface 84 and an outlet 94 located at the suction side surface 86 .
- the inlet 92 receives a portion P of core airflow F. The portion P is communicated through the flow channel 82 prior to being expelled from the outlet 94 .
- the inlet 92 and the outlet 94 may be located anywhere on the pressure side surface 84 and the suction side surface 86 , respectively.
- the flow channel 82 may define any size or shape. In one non-limiting embodiment, the flow channel 82 extends along a curved path. However, the flow channel 82 could follow a linear or non-linear path, a curved path, or any other configuration within the scope of this disclosure. In addition, although shown with a single flow channel, the airfoil 180 could include multiple flow channels.
- FIG. 4 illustrates another exemplary flow channel 182 of an airfoil 180 . Only a section of the airfoil 180 is illustrated in FIG. 4 .
- like reference numerals designate like elements where appropriate and reference numerals with the addition of (100) or multiples thereof designate modified elements that are understood to incorporate the same features and benefits of the corresponding original elements.
- the flow channel 182 extends through an interior of the airfoil 180 between a pressure side surface 184 and a suction side surface 186 and includes a radial dimension RD.
- An inlet 192 of the flow channel 182 may be positioned at a first radial location R 1 of the span of the airfoil 180 , and an outlet 194 may exit the suction side surface 186 at a second radial location R 2 of the span that is different from the first radial location R 1 .
- the radial dimension RD of the flow channel 182 includes a span that is generally less than the total span of the airfoil 180 .
- FIG. 5 illustrates yet another airfoil 280 having a flow channel 282 .
- an outlet 294 of the flow channel 282 exits the suction side surface 286 at a position that is upstream from a throat area TA that extends between the airfoil 280 and a circumferentially adjacent airfoil 280 - 2 .
- the throat area TA is a planar opening with a periphery bounded in a radial direction by an outer diameter platform (or casing) and an inner diameter platform and peripherally bounded in the circumferential direction by the suction side surface 286 of the airfoil 280 and a trailing edge 290 of the adjacent airfoil 280 - 2 .
- the outlet 294 of the flow channel 282 could exit the suction side surface 286 of the airfoil 280 at a position that is downstream from the throat area TA. This is schematically shown in FIG. 6 . It should be understood that the outlet 294 of the flow channel 282 could exit at any location of the suction side surface 286 and that an inlet 292 could be located anywhere along the pressure side surface 284 .
- FIGS. 7A and 7B illustrate embodiments of vanes 166 that include flow channels 182 A similar to those described above.
- the vanes 166 include an airfoil 180 A having a flow channel 182 A that extends between an inlet 192 A at a pressure side surface 184 A and an outlet 194 A at a suction side surface 186 A.
- the flow channel 182 A defines a flow passage that extends from a tip 96 A to a root 98 A of the airfoil 180 A.
- the flow channel 182 A transcends in both chordwise and spanwise directions of the airfoil 180 A.
- the flow channel 182 A could define a flow passage that extends from the root 98 A toward the tip 96 A of the airfoil 180 A.
- the inlet 192 A is still positioned along the pressure side surface 184 A and the outlet 194 A still exits at the suction side surface 186 A of the airfoil 180 A.
- FIGS. 8A and 8B illustrate embodiments of blades 176 that include flow channels 182 B similar to those described above.
- the blades 176 include an airfoil 180 B having a flow channel 182 B that extends between an inlet 192 B at a pressure side surface 184 B and an outlet 194 B at a suction side surface 186 B.
- the flow channel 182 B defines a flow passage that extends from a tip 96 B to a root 98 B of the airfoil 180 B.
- the flow channel 182 B transcends in both chordwise and spanwise directions of the airfoil 180 B.
- the flow channel 182 B could define a flow passage that extends from the root 98 B to the tip 96 B of the airfoil 180 B.
- the inlet 192 B is still positioned along the pressure side surface 184 B and the outlet 194 B exits at the suction side surface 186 B of the airfoil 180 B.
- FIGS. 9A, 9B and 9C illustrate multiple configurations of a variable area section 200 having a variable vane assembly 264 and a rotor assembly 272 downstream from the variable vane assembly 264 .
- the variable area section 200 could include additional stages of vane and rotor assemblies within the scope of this disclosure.
- the variable vane assembly 264 includes a plurality of variable vanes 266 (two shown) that are rotatable between open ( FIG. 9A ), nominal ( FIG. 9B ), and closed ( FIG. 9C ) positions in order to vary an amount of core airflow F that is communicated through cascade passages 265 that extend between adjacent vanes 266 of the variable vane assembly 264 to impinge upon the downstream rotor assembly 272 .
- the rotor assembly 272 includes a plurality of blades 276 (two shown). The blades 276 rotate about the engine central longitudinal axis A (see FIG. 1 ) to extract energy from the core airflow F.
- variable vanes 266 include flow channels 282 A and the blades 276 include flow channels 282 B.
- a portion P of the core airflow F may be communicated through the flow channels 282 A in order to alter the flow stagnation point of the downstream blades 276 .
- a second portion P 2 of the core airflow F may be communicated through the flow channels 282 B in order to alter the flow stagnation point on components downstream of the blades 276 . In this way, any negative effects caused by incidence angle variation can be substantially ameliorated.
- variable vanes 266 include flow channels 282 A and the blades 276 include flow channels 282 B.
- a portion P of the core airflow F may be communicated through the flow channels 282 A in order to alter the flow stagnation point of the variable vanes 266 .
- a second portion P 2 of the core airflow F may be communicated through the flow channels 282 B in order to alter the flow stagnation point of the blades 276 . In this way, any negative effects caused by incidence angle variation can be substantially ameliorated.
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Abstract
Description
- This disclosure relates to a gas turbine engine, and more particularly to a gas turbine engine airfoil having an auxiliary flow channel for receiving and communicating a portion of core airflow through the airfoil.
- Gas turbine engines typically include at least a compressor section, a combustor section, and a turbine section. In general, during operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases flow through the turbine section, which extracts energy from the hot combustion gases to power the compressor section as well as other gas turbine engine loads.
- The compressor and turbine sections typically include alternating rows of rotating blades and flow directing vanes. In the turbine section, the rotating blades extract energy from the core airflow that is communicated through the gas turbine engine, while the vanes direct the core airflow to a downstream row of blades.
- The vanes can be manufactured to a fixed flow area that is optimized for a single flight point. Alternatively, it is possible to alter the flow area (i.e., cascade channel) between two adjacent vanes by providing one or more variable vanes that rotate about a given axis. Altering the flow area in this manner can expose downstream components to incidence angle variation. For example, rotating the variable vanes may alter the incidence angle at which hot combustion gases impinge upon rotor blades located downstream from the variable vanes, thereby potentially moving the flow stagnation point to a non-optimal location.
- A component for a gas turbine engine, according to an exemplary aspect of the present disclosure includes, among other things, an airfoil that includes a pressure side surface and a suction side surface that join together at a leading edge and a trailing edge and a flow channel that extends between the pressure side surface and the suction side surface.
- In a further non-limiting embodiment of the foregoing component, the component is a vane or a blade.
- In a further non-limiting embodiment of either of the foregoing components, the component is a mid-turbine frame, an exit guide vane or a fan blade.
- In a further non-limiting embodiment of any of the foregoing components, the flow channel includes an inlet at the pressure side surface and an outlet at the suction side surface.
- In a further non-limiting embodiment of any of the foregoing components, the flow channel includes a radial dimension that is less than a total span of the airfoil.
- In a further non-limiting embodiment of any of the foregoing components, the flow channel includes an outlet positioned upstream from a throat area that extends between the airfoil and an adjacent airfoil.
- In a further non-limiting embodiment of any of the foregoing components, the flow channel includes an outlet positioned downstream from a throat area that extends between the airfoil and an adjacent airfoil.
- In a further non-limiting embodiment of any of the foregoing components, the flow channel defines a flow passage that extends from a tip toward a root of the airfoil.
- In a further non-limiting embodiment of any of the foregoing components, the flow channel defines a flow passage that extends from a root toward a tip of the airfoil.
- In a further non-limiting embodiment of any of the foregoing components, the flow channel extends along a curved path.
- A gas turbine engine, according to another exemplary aspect of the present disclosure includes, among other things, a variable area section, a variable vane disposed within the variable area section and a blade downstream from the variable vane. At least one of the variable vane and the blade includes a flow channel having an inlet at a pressure side surface and an outlet at a suction side surface of an airfoil of either the variable vane or the blade.
- In a further non-limiting embodiment of the foregoing gas turbine engine, the variable area section is a variable area turbine section.
- In a further non-limiting embodiment of either of the foregoing gas turbine engines, the variable area section is a variable area compressor section.
- In a further non-limiting embodiment of any of the foregoing gas turbine engines, both the variable vane and the blade include the flow channel.
- In a further non-limiting embodiment of any of the foregoing gas turbine engines, the flow channel extends in each of a chordwise and a spanwise direction of the airfoil.
- A method of operating a gas turbine engine according to another exemplary aspect of the present disclosure includes, among other things, communicating a portion of core airflow through a flow channel that extends between a pressure side surface and a suction side surface of at least one airfoil of the gas turbine engine.
- In a further non-limiting embodiment of the foregoing method, the method includes the step of rotating a variable vane prior to the step of communicating.
- In a further non-limiting embodiment of either of the foregoing methods, the method includes communicating the portion of airflow into an inlet of the flow channel located along the pressure side surface and expelling the portion of airflow through an outlet of the flow channel located along the suction side surface.
- In a further non-limiting embodiment of any of the foregoing methods, the method includes the step of influencing incidence angle variation of a downstream component with the portion of core airflow.
- In a further non-limiting embodiment of any of the foregoing methods, the method includes the step of communicating a second portion of core airflow through a flow channel of the downstream component.
- The embodiments, examples and alternatives of the preceding paragraphs, the claims, or the following description and drawings, including any of their various aspects or respective individual features, may be taken independently or in any combination. Features described in connection with one embodiment are applicable to all embodiments, unless such features are incompatible.
- The various features and advantages of this disclosure will become apparent to those skilled in the art from the following detailed description. The drawings that accompany the detailed description can be briefly described as follows.
-
FIG. 1 illustrates a schematic, cross-sectional view of a gas turbine engine. -
FIG. 2 illustrates a variable area section of a gas turbine engine. -
FIG. 3 illustrates an airfoil having a flow channel. -
FIG. 4 illustrates another airfoil flow channel. -
FIG. 5 illustrates another airfoil flow channel. -
FIG. 6 illustrates yet another airfoil flow channel. -
FIGS. 7A and 7B illustrate vanes having flow channels. -
FIGS. 8A and 8B illustrate blades having flow channels. -
FIGS. 9A, 9B and 9C illustrate multiple configurations of a variable area gas turbine engine section that includes one or more components having flow channels. - This disclosure is directed to gas turbine engine components that include auxiliary flow channels. A variable area section of a gas turbine engine may employ one or more airfoils that include a flow channel extending between a pressure side surface and a suction side surface. Core airflow may be communicated through the flow channel to influence a location of the flow stagnation point on the airfoil and downstream components. For example, a portion of core airflow may be communicated through the flow channel in order to modify the incidence angle of the core flow on the airfoil. In another example, a portion of core airflow may be communicated through the flow channel in order to modify the incidence angle at which core airflow impinges on downstream hardware.
-
FIG. 1 schematically illustrates agas turbine engine 20. Thegas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates afan section 22, acompressor section 24, acombustor section 26 and aturbine section 28. Alternative engines might include an augmentor section (not shown) among other systems or features. Thefan section 22 drives air along a bypass flow path B in a bypass duct defined within anacelle 15, while thecompressor section 24 drives air along a core flow path C for compression and communication into thecombustor section 26 then expansion through theturbine section 28. Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. - The
exemplary engine 20 generally includes alow speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an enginestatic structure 36 viaseveral bearing systems 38. It should be understood thatvarious bearing systems 38 at various locations may alternatively or additionally be provided, and the location ofbearing systems 38 may be varied as appropriate to the application. - The
low speed spool 30 generally includes aninner shaft 40 that interconnects afan 42, a first (or low) pressure compressor 44 and a first (or low)pressure turbine 46. Theinner shaft 40 is connected to thefan 42 through a speed change mechanism, which in exemplarygas turbine engine 20 is illustrated as a gearedarchitecture 48 to drive thefan 42 at a lower speed than thelow speed spool 30. Thehigh speed spool 32 includes anouter shaft 50 that interconnects a second (or high)pressure compressor 52 and a second (or high)pressure turbine 54. Acombustor 56 is arranged inexemplary gas turbine 20 between thehigh pressure compressor 52 and thehigh pressure turbine 54. Amid-turbine frame 57 of the enginestatic structure 36 is arranged generally between thehigh pressure turbine 54 and thelow pressure turbine 46. Themid-turbine frame 57 furthersupports bearing systems 38 in theturbine section 28. Theinner shaft 40 and theouter shaft 50 are concentric and rotate via bearingsystems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes. - The core airflow is compressed by the low pressure compressor 44 then the
high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over thehigh pressure turbine 54 andlow pressure turbine 46. Themid-turbine frame 57 includesairfoils 59 which are in the core airflow path C. Theturbines low speed spool 30 andhigh speed spool 32 in response to the expansion. It will be appreciated that each of the positions of thefan section 22,compressor section 24,combustor section 26,turbine section 28, and fandrive gear system 48 may be varied. For example,gear system 48 may be located aft ofcombustor section 26 or even aft ofturbine section 28, andfan section 22 may be positioned forward or aft of the location ofgear system 48. - The
engine 20 in one example is a high-bypass geared aircraft engine. In a further example, theengine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the gearedarchitecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and thelow pressure turbine 46 has a pressure ratio that is greater than about five. In one disclosed embodiment, theengine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and thelow pressure turbine 46 has a pressure ratio that is greater than about five 5:1.Low pressure turbine 46 pressure ratio is pressure measured prior to inlet oflow pressure turbine 46 as related to the pressure at the outlet of thelow pressure turbine 46 prior to an exhaust nozzle. The gearedarchitecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans. - A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The
fan section 22 of theengine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)]0.5. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second. -
FIG. 2 illustrates avariable area section 100 that may be incorporated into a gas turbine engine, such as thegas turbine engine 20 ofFIG. 1 . In one embodiment, thevariable area section 100 is a variable area turbine section. In another embodiment, thevariable area section 100 is a variable area compressor section. Thevariable area section 100 may make up part of or the entirety of thecompressor section 24 and/or theturbine section 28 of thegas turbine engine 20. In other words, these sections could have components in addition to what is shown inFIG. 2 . - The
variable area section 100 includes one or more stages of alternating rows of vanes and blades. In the illustrated embodiment, thevariable area section 100 includes a single stage (i.e., one row of vanes and one row of blades); however, additional stages could be incorporated into thevariable area section 100. - A
case structure 62 houses the components of thevariable area section 100. In one embodiment, thecase structure 62 is an outer engine casing that circumscribes the components of thevariable area section 100. - In one non-limiting embodiment, the
variable area section 100 includes avariable vane assembly 64 having at least onevariable vane 66. Thevariable vane assembly 64 could include an array ofvariable vanes 66 circumferentially disposed about the engine centerline longitudinal axis A. Alternatively, thevariable vane assembly 64 could include a combination of both fixed and variable vanes. - The
variable vanes 66 are selectively configurable to change a flow parameter associated with thevariable area section 100. For example, eachvariable vane 66 may be rotated or pivoted (via an actuation system) about a spindle axis SA in order to change the rotational positioning of thevariable vane 66. Rotating the variable vane(s) 66 changes the amount of core airflow F that enters thevariable area section 100, thereby influencing the flow area of thevariable area section 100. - The variable vane(s) 66 is rotatable relative to an
inner platform 68 and anouter platform 70 of thevariable vane assembly 64. Theinner platform 68 and theouter platform 70 may be mounted to thecase structure 62 in any known manner. - A
rotor assembly 72 is positioned downstream from thevariable vane assembly 64. Therotor assembly 72 includes at least onerotor disk 74 that carries one ormore rotor blades 76. Therotor blades 76 rotate about the engine central longitudinal axis A to extract energy from the core airflow F, thereby moving therotor disk 74 and powering various gas turbine engine loads. - The
rotor blades 76 rotate relative to blade outer air seals (BOAS) 78 that establish a radially outer flow path boundary for channeling the core airflow F through thevariable area section 100. TheBOAS 78 may mount to thecase structure 62 and extend in relationship to a tip of eachrotating blade 76 in order to seal between theblades 76 and thecase structure 62. - Altering the flow area associated with the
variable area section 100 by moving the variable vane(s) 66 can change the incidence angle α at which core airflow F impinges upon therotor blades 76 of thedownstream rotor assembly 72. Incidence angle variation can negatively influence gas turbine engine efficiency by altering the stagnation point on the airfoil. Accordingly, thevariable vanes 66 and/orrotor blades 76 may include airfoils having auxiliary flow channels for addressing such variations. Airfoils of this type are discussed in greater detail below with respect toFIGS. 3-9 . -
FIG. 3 illustrates a cross-sectional view of anexemplary airfoil 80 that includes aflow channel 82. Theairfoil 80 may be part of a vane and/or a blade. For example, theairfoil 80 could be a section of avariable vane 66 of thevariable vane assembly 64 and/or arotor blade 76 of therotor assembly 72 ofFIG. 2 (see, for example,FIGS. 7A, 7B and 8A, 8B ). Theairfoil 80 could also be part of other gas turbine engine components including but not limited to the mid-turbine frame, a fan exit guide vane or a fan blade. - In one embodiment, the
flow channel 82 extends inside of theairfoil 80 between apressure side surface 84 and asuction side surface 86 of theairfoil 80. Thepressure side surface 84 and thesuction side surface 86 are spaced apart from one another and generally join together at aleading edge 88 and a trailingedge 90 of theairfoil 80. - In one embodiment, the
flow channel 82 includes aninlet 92 located at thepressure side surface 84 and anoutlet 94 located at thesuction side surface 86. Theinlet 92 receives a portion P of core airflow F. The portion P is communicated through theflow channel 82 prior to being expelled from theoutlet 94. Theinlet 92 and theoutlet 94 may be located anywhere on thepressure side surface 84 and thesuction side surface 86, respectively. - The
flow channel 82 may define any size or shape. In one non-limiting embodiment, theflow channel 82 extends along a curved path. However, theflow channel 82 could follow a linear or non-linear path, a curved path, or any other configuration within the scope of this disclosure. In addition, although shown with a single flow channel, theairfoil 180 could include multiple flow channels. -
FIG. 4 illustrates anotherexemplary flow channel 182 of anairfoil 180. Only a section of theairfoil 180 is illustrated inFIG. 4 . In this disclosure, like reference numerals designate like elements where appropriate and reference numerals with the addition of (100) or multiples thereof designate modified elements that are understood to incorporate the same features and benefits of the corresponding original elements. - In this embodiment, the
flow channel 182 extends through an interior of theairfoil 180 between apressure side surface 184 and asuction side surface 186 and includes a radial dimension RD. Aninlet 192 of theflow channel 182 may be positioned at a first radial location R1 of the span of theairfoil 180, and anoutlet 194 may exit thesuction side surface 186 at a second radial location R2 of the span that is different from the first radial location R1. In one non-limiting embodiment, the radial dimension RD of theflow channel 182 includes a span that is generally less than the total span of theairfoil 180. -
FIG. 5 illustrates yet anotherairfoil 280 having aflow channel 282. In one embodiment, anoutlet 294 of theflow channel 282 exits thesuction side surface 286 at a position that is upstream from a throat area TA that extends between theairfoil 280 and a circumferentially adjacent airfoil 280-2. The throat area TA is a planar opening with a periphery bounded in a radial direction by an outer diameter platform (or casing) and an inner diameter platform and peripherally bounded in the circumferential direction by thesuction side surface 286 of theairfoil 280 and a trailingedge 290 of the adjacent airfoil 280-2. - Alternatively, the
outlet 294 of theflow channel 282 could exit thesuction side surface 286 of theairfoil 280 at a position that is downstream from the throat area TA. This is schematically shown inFIG. 6 . It should be understood that theoutlet 294 of theflow channel 282 could exit at any location of thesuction side surface 286 and that aninlet 292 could be located anywhere along thepressure side surface 284. -
FIGS. 7A and 7B illustrate embodiments ofvanes 166 that includeflow channels 182A similar to those described above. Thevanes 166 include anairfoil 180A having aflow channel 182A that extends between aninlet 192A at apressure side surface 184A and anoutlet 194A at asuction side surface 186A. Referring toFIG. 7A , theflow channel 182A defines a flow passage that extends from atip 96A to aroot 98A of theairfoil 180A. In other words, theflow channel 182A transcends in both chordwise and spanwise directions of theairfoil 180A. - Alternatively, as shown in
FIG. 7B , theflow channel 182A could define a flow passage that extends from theroot 98A toward thetip 96A of theairfoil 180A. In this embodiment, theinlet 192A is still positioned along the pressure side surface 184A and theoutlet 194A still exits at thesuction side surface 186A of theairfoil 180A. -
FIGS. 8A and 8B illustrate embodiments ofblades 176 that includeflow channels 182B similar to those described above. Theblades 176 include anairfoil 180B having aflow channel 182B that extends between aninlet 192B at apressure side surface 184B and anoutlet 194B at asuction side surface 186B. Referring toFIG. 8A , theflow channel 182B defines a flow passage that extends from atip 96B to aroot 98B of theairfoil 180B. In other words, theflow channel 182B transcends in both chordwise and spanwise directions of theairfoil 180B. - Alternatively, as shown in
FIG. 8B , theflow channel 182B could define a flow passage that extends from theroot 98B to thetip 96B of theairfoil 180B. In this embodiment, theinlet 192B is still positioned along thepressure side surface 184B and theoutlet 194B exits at thesuction side surface 186B of theairfoil 180B. -
FIGS. 9A, 9B and 9C illustrate multiple configurations of avariable area section 200 having avariable vane assembly 264 and arotor assembly 272 downstream from thevariable vane assembly 264. Thevariable area section 200 could include additional stages of vane and rotor assemblies within the scope of this disclosure. - The
variable vane assembly 264 includes a plurality of variable vanes 266 (two shown) that are rotatable between open (FIG. 9A ), nominal (FIG. 9B ), and closed (FIG. 9C ) positions in order to vary an amount of core airflow F that is communicated throughcascade passages 265 that extend betweenadjacent vanes 266 of thevariable vane assembly 264 to impinge upon thedownstream rotor assembly 272. Therotor assembly 272 includes a plurality of blades 276 (two shown). Theblades 276 rotate about the engine central longitudinal axis A (seeFIG. 1 ) to extract energy from the core airflow F. - In one embodiment, the
variable vanes 266 includeflow channels 282A and theblades 276 includeflow channels 282B. A portion P of the core airflow F may be communicated through theflow channels 282A in order to alter the flow stagnation point of thedownstream blades 276. A second portion P2 of the core airflow F may be communicated through theflow channels 282B in order to alter the flow stagnation point on components downstream of theblades 276. In this way, any negative effects caused by incidence angle variation can be substantially ameliorated. - In the same embodiment, the
variable vanes 266 includeflow channels 282A and theblades 276 includeflow channels 282B. A portion P of the core airflow F may be communicated through theflow channels 282A in order to alter the flow stagnation point of thevariable vanes 266. A second portion P2 of the core airflow F may be communicated through theflow channels 282B in order to alter the flow stagnation point of theblades 276. In this way, any negative effects caused by incidence angle variation can be substantially ameliorated. - Although the different non-limiting embodiments are illustrated as having specific components, the embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from any of the non-limiting embodiments in combination with features or components from any of the other non-limiting embodiments.
- It should be understood that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be understood that although a particular component arrangement is disclosed and illustrated in these exemplary embodiments, other arrangements could also benefit from the teachings of this disclosure.
- The foregoing description shall be interpreted as illustrative and not in any limiting sense. A worker of ordinary skill in the art would understand that certain modifications could come within the scope of this disclosure. For these reasons, the following claims should be studied to determine the true scope and content of this disclosure.
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FR2438156A1 (en) * | 1978-10-05 | 1980-04-30 | Alsthom Atlantique | Turbine blade ring - has holes drilled in blades to relieve fluid build up due to centrifugal force |
US6203269B1 (en) * | 1999-02-25 | 2001-03-20 | United Technologies Corporation | Centrifugal air flow control |
US6435815B2 (en) * | 2000-01-22 | 2002-08-20 | Rolls-Royce Plc | Aerofoil for an axial flow turbo machine |
US7223066B2 (en) * | 2003-05-27 | 2007-05-29 | Rolls-Royce Plc | Variable vane arrangement for a turbomachine |
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US10151322B2 (en) * | 2016-05-20 | 2018-12-11 | United Technologies Corporation | Tandem tip blade |
US20180195528A1 (en) * | 2017-01-09 | 2018-07-12 | Rolls-Royce Coporation | Fluid diodes with ridges to control boundary layer in axial compressor stator vane |
US10519976B2 (en) * | 2017-01-09 | 2019-12-31 | Rolls-Royce Corporation | Fluid diodes with ridges to control boundary layer in axial compressor stator vane |
CN107514292A (en) * | 2017-09-30 | 2017-12-26 | 南京赛达机械制造有限公司 | A kind of torsion fracture resistant turbine blade |
US20190301301A1 (en) * | 2018-04-02 | 2019-10-03 | General Electric Company | Cooling structure for a turbomachinery component |
US10808572B2 (en) * | 2018-04-02 | 2020-10-20 | General Electric Company | Cooling structure for a turbomachinery component |
JP2020139421A (en) * | 2019-02-27 | 2020-09-03 | 三菱重工業株式会社 | Blade and rotating machine comprising the same |
US11391159B2 (en) * | 2019-02-27 | 2022-07-19 | Mitsubishi Heavy Industries, Ltd. | Blade and rotary machine having the same |
JP7221078B2 (en) | 2019-02-27 | 2023-02-13 | 三菱重工業株式会社 | Wings and rotating machines equipped with them [2006.01] |
US20240102395A1 (en) * | 2022-09-27 | 2024-03-28 | Pratt & Whitney Canada Corp. | Stator vane for a gas turbine engine |
Also Published As
Publication number | Publication date |
---|---|
EP3063374A1 (en) | 2016-09-07 |
WO2015065659A1 (en) | 2015-05-07 |
EP3063374A4 (en) | 2017-08-09 |
US10280757B2 (en) | 2019-05-07 |
EP3063374B1 (en) | 2023-07-19 |
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